Gas turbine engine with geared turbofan and oil thermal management system

ABSTRACT

A gas turbine engine includes a fan, a compressor section, a combustion section, and a turbine section. A fan drive gear system is configured for driving the fan at a speed different than the turbine section. A lubricant system includes a lubricant pump delivering lubricant to an outlet line. The outlet line splits into at least a hot line and into a cool line. The hot line is directed primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant. The cool line is directed through one or more heat exchangers at which the lubricant is cooled, and the cool line then is routed to the fan drive gear system. At least one of the one or more heat exchangers is a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section. The fuel/oil cooler is downstream of a point where the outlet line splits into the at least the hot line and the cool line, such that the hot line is not directed through the fuel/oil cooler. A method is also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of co-pending application Ser. No. 13/794,976 filed Mar. 12, 2013 that is a continuation of co-pending application Ser. No. 13/361,997 filed Jan. 31, 2012, now U.S. Pat. No. 8,443,582 granted May 21, 2013.

BACKGROUND

This application relates to an oil system for providing oil to a gear associated with a geared turbofan in a gas turbine engine.

Gas turbine engines are known, and typically include a fan delivering air into a compressor section. Compressed air from the compressor section is delivered into a combustion section, mixed with fuel, and ignited. Products of this combustion pass downstream over turbine rotors which are driven to rotate.

A low pressure turbine rotor drives a low pressure compressor, and traditionally has driven a fan at the same rate of speed.

More recently, a gear reduction has been included between the low pressure turbine and the fan such that the fan and the low pressure compressor can rotate at different speeds.

Oil management systems are known, and typically provide oil to engine bearings and other locations within the engine. As a result of gears being added to turbofan engines, additional components require cooling, thereby necessitating new cooling systems and methodologies.

SUMMARY

A gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan, a compressor section, a combustion section, and a turbine section. A fan drive gear system is configured for driving the fan at a speed different than the turbine section. A lubricant system includes a lubricant pump delivering lubricant to an outlet line. The outlet line splits into at least a hot line and into a cool line. The hot line is directed primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant. The cool line is directed through one or more heat exchangers at which the lubricant is cooled, and the cool line then is routed to the fan drive gear system. At least one of the one or more heat exchangers is a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section. The fuel/oil cooler is downstream of a point where the outlet line splits into the at least the hot line and the cool line, such that the hot line is not directed through the fuel/oil cooler.

In a further embodiment of any of the foregoing gas turbine engines, the locations in the engine that are not intended to receive cooler lubricant include bearings associated with at least the turbine section.

In a further embodiment of any of the foregoing gas turbine engines, the cool line supplies lubricant to a bearing in the fan drive gear system.

In a further embodiment of any of the foregoing gas turbine engines, the bearing in the fan drive gear system includes a journal bearing.

In a further embodiment of any of the foregoing gas turbine engines, a valve selectively supplies lubricant from the hot line to the fan drive gear system when additional lubricant is necessary.

In a further embodiment of any of the foregoing gas turbine engines, a valve is positioned on the cool line, at a location downstream of a point where the hot line splits off from the cool line, and the valve splits the cool line into two lines, with a first line being directed through the one or more heat exchanger, and a second line being directed through at least one other cooler.

In a further embodiment of any of the foregoing gas turbine engines, the lubricant in the hot line does not intermix back into the cool line until a point after the cool line has been routed to the bearing for the fan drive gear system.

In a further embodiment of any of the foregoing gas turbine engines, the lubricant in the hot line is intermixed with a portion of the lubricant from the cool line prior to being delivered to the locations in the gas turbine engine that are not intended to receive cooler lubricant.

In a further embodiment of any of the foregoing gas turbine engines, the compressor section includes a first compressor section axially forward of a second compressor section.

In a further embodiment of any of the foregoing gas turbine engines, the turbine section includes a fan drive turbine and a second turbine. The fan drive turbine is configured for driving the first compressor section, and the fan drive gear system such that the fan and the first compressor are driven at different speeds.

In a further embodiment of any of the foregoing gas turbine engines, the fan drive gear system is configured to provide a speed reduction between the fan and the fan drive turbine greater than about 2.3:1.

In a further embodiment of any of the foregoing gas turbine engines, the fan drive gear system is configured to provide a speed reduction between the fan and the fan drive turbine greater than about 2.5:1.

In a further embodiment of any of the foregoing gas turbine engines, the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio greater than about six (6).

In a further embodiment of any of the foregoing gas turbine engines, the bypass ratio is greater than about ten (10).

In a further embodiment of any of the foregoing gas turbine engines, the gas turbine engine includes a fan pressure ratio of less than about 1.45.

In a further embodiment of any of the foregoing gas turbine engines, the fan is configured to rotate with a fan tip speed that is less than about 1150 ft/second.

In a further embodiment of any of the foregoing gas turbine engines, the fan drive turbine is configured to operate at a pressure ratio that is greater than about 5:1.

In a further embodiment of any of the foregoing gas turbine engines, the second turbine section operates at a higher pressure than the fan drive turbine and is configured to drive the second compressor section.

A method of designing a gas turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes defining a fan configured for rotation about an axis. A compressor section, a combustor section, and a turbine section are configured to define a core flow path. A fan drive gear system is configured for driving the fan at a speed different than the turbine section. A lubricant system is defined to include a lubricant pump for delivering lubricant to an outlet line. The outlet is defined to split into at least a hot line and into a cool line. The hot line is configured for directing lubricant primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant. The cool line is configured for directing lubricant through one or more heat exchangers for cooling lubricant. The cool line is configured for routing to the fan drive gear system. At least one of the one or more heat exchangers is configured as a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section. The fuel/oil cooler is downstream of a point where the outlet line splits into the at least the hot line and the cool line, such that the hot line is not directed through the fuel/oil cooler.

In a further embodiment of any of the foregoing methods, includes configuring the locations in the engine that are not intended to receive cooler lubricant to include bearings associated with at least the turbine section.

In a further embodiment of any of the foregoing methods, the cool line is configured to supply lubricant to a journal bearing in the fan drive gear system.

In a further embodiment of any of the foregoing methods, includes defining a valve for selectively supplying lubricant from the hot line to the fan drive gear system when additional lubricant is necessary.

In a further embodiment of any of the foregoing methods, includes configuring a valve to be positioned on the cool line, at a location downstream of a point where the hot line splits off from the cool line, and configuring the valve for splitting the cool line into two lines, with a first line being directed through the one or more heat exchanger, and a second line being directed through at least one other cooler.

In a further embodiment of any of the foregoing methods, includes defining the system such that lubricant in the hot line does not intermix back into said cool line until a point after the cool line has been routed to a bearing for the fan drive gear system.

In a further embodiment of any of the foregoing methods, includes defining the system such that lubricant in the hot line is intermixed with a portion of the lubricant from the cool line prior to being delivered to the locations in the gas turbine engine that are not intended to receive cooler lubricant.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 is a schematic of an oil management system for the gas turbine engine of FIG. 1.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath B while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 (shown schematically) to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the previously mentioned expansion.

The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7 ° R)]^(0.5). The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

FIG. 2 is an oil management system for the gas turbine engine system of FIG. 1. The oil management system 140 is utilized in association with a fuel system 143, and a variable frequency generator 160 and its oil cooling system circuit 162.

Fuel from a fuel tank 142 passes to a fuel/oil cooler 144. The fuel is heated, and cools a lubricant, as will be explained below. A main fuel pump 146 drives the fuel into further fuel lines 243 and then into nozzles 148 in a combustor, such as combustor 26 as shown in FIG. 1. It is known in the art to heat the fuel to improve the efficiency of the overall engine. The fuel/oil cooler 144 provides this function.

At the same time, the variable frequency generator 160 is driven by turbine rotors to generate electricity for various uses on an aircraft. As shown in oil cooling system circuit 162, the oil passes through an oil-to-oil cooler 164, and may also pass through an air-to-oil cooler 66 before returning to the variable frequency generator 160.

An oil supply system 150 includes a main oil pump 70 taking oil from a main oil tank 72. The terms “oil” and “lubricant” are used interchangeably in this application and cover a fluid used to lubricate surfaces subject to relative rotation. The oil is delivered through a downstream line 73, and split between two lines 74 and 75. Line 74 is sent directly to line 86 without cooling. A modulating valve 76 is controlled to achieve a desired fuel temperature. As an example, a sensor 300 may send a signal to a control regarding a sensed temperature of the fuel downstream of the fuel oil cooler 144. The valve 76 routes the volume of oil between line 78 and 80 to achieve the desired temperature of the fuel.

The oil passing to line 78 passes through the fuel/oil cooler 144 and heats the fuel. The oil is cooled before returning to a common downstream line 82. The downstream line 82 could be called a “cool” oil line, as the oil will be cooler than the oil in “hot” line 74 which has not been cooled in any heat exchanger. For purposes of this application, line 75 is seen as part of the “cool” line even though the lubricant has yet to be cooled.

The oil directed by the valve 76 into line 80 passes through an air-to-oil cooler at 68 which is exposed to air which is cooler than the oil in line 80, and which cools the oil. Downstream of the air-to-oil cooler 68, the oil passes through the oil-to-oil cooler 164, and may actually be somewhat heated by cooling the oil for the variable frequency generator. Still, the oil reaching line 82 downstream of the oil-to-oil cooler 164 will be significantly cooler than the oil in line 74. Some of the oil in line 82 is directed into a line 84, to a journal bearing 152, and to the fan drive gears 154. Thus, cooler oil is supplied to the bearing 152 and gears 154 than is supplied from the line 74. As can be seen, a line 86 branches off of the “cool” line 82 at or near the point at which “cool” line 84 breaks away to go to the journal bearing 152 and the gears 154. A return line 88 is downstream of the journal bearing 152 and gears 154. The lubricant in line 86 mixes with the lubricant in “hot” line 74, but downstream of the branch line 84.

It is desirable to provide cooler oil to these locations than is necessary to be supplied to bearings 90, or other locations associated with the engine. The bearings 90 as shown in FIG. 2 may equate to the several locations of bearings 38 as shown in FIG. 1. The journal bearing 152 and the fan drive gears 154 would be part of the geared architecture 48 as shown in FIG. 1.

On the other hand, cooling all of the oil associated with the engine bearings 90 would reduce the overall efficiency of the engine. Thus, splitting the oil, and cooling the oil to be directed to the bearings 152 and/or gears 154 provides cooler oil to those locations, while still allowing the hotter oil to be directed to locations that do not need cooler oil.

In addition, a valve 92 can selectively direct additional oil to the gears 154 if additional oil is necessary, such as at high power times. At other times, the valve 92 may direct lubricant through line 94 back to a return line 95 leading back to the oil tank 72.

The overall configuration thus results in an oil supply system which directs hotter oil to the locations which do not need cooler oil, but which also cools oil to be directed to areas associated with the fan drive gear.

Further details of a similar oil management system are disclosed in co-pending U.S. patent application Ser. No. 13/362,094, entitled “Gas Turbine Engine With Geared Turbofan and Oil Thermal Management System With Unique Heat Exchanger Structure, owned by the assignee of the present application.

Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention. 

What is claimed is:
 1. A gas turbine engine comprising: a fan; a compressor section; a combustion section; a turbine section; a fan drive gear system configured for driving the fan at a speed different than the turbine section; a lubricant system including a lubricant pump delivering lubricant to an outlet line, said outlet line splitting into at least a hot line and into a cool line, said hot line being directed primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant, and said cool line being directed through one or more heat exchangers at which the lubricant is cooled, and said cool line then being routed to said fan drive gear system; and at least one of said one or more heat exchangers is a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section, and said fuel/oil cooler being downstream of a point where said outlet line splits into said at least said hot line and said cool line, such that said hot line is not directed through said fuel/oil cooler.
 2. The gas turbine engine as set forth in claim 1, wherein said locations in the engine that are not intended to receive cooler lubricant include bearings associated with at least the turbine section.
 3. The gas turbine engine as set forth in claim 2, wherein said cool line supplies lubricant to a bearing in said fan drive gear system.
 4. The gas turbine engine as set forth in claim 3, wherein the bearing in said fan drive gear system comprises a journal bearing.
 5. The gas turbine engine as set forth in claim 3, wherein a valve selectively supplies lubricant from said hot line to the fan drive gear system when additional lubricant is necessary.
 6. The gas turbine engine as set forth in claim 3, wherein a valve is positioned on said cool line, at a location downstream of a point where said hot line splits off from said cool line, and said valve splitting said cool line into two lines, with a first line being directed through said one or more heat exchanger, and a second line being directed through at least one other cooler.
 7. The gas turbine engine as set forth in claim 5, wherein lubricant in said hot line does not intermix back into said cool line until a point after said cool line has been routed to the bearing for said fan drive gear system.
 8. The gas turbine engine as set forth in claim 5, wherein the lubricant in the hot line is intermixed with a portion of the lubricant from the cool line prior to being delivered to said locations in the gas turbine engine that are not intended to receive cooler lubricant.
 9. The gas turbine engine as set forth in claim 8, wherein the compressor section includes a first compressor section axially forward of a second compressor section.
 10. The gas turbine engine as set forth in claim 9, wherein the turbine section includes a fan drive turbine and a second turbine, and said fan drive turbine configured for driving said first compressor section, and the fan drive gear system such that said fan and said first compressor are driven at different speeds.
 11. The gas turbine engine as set forth in claim 10, wherein the fan drive gear system is configured to provide a speed reduction between the fan and the fan drive turbine greater than about 2.3:1.
 12. The gas turbine engine as set forth in claim 11, wherein the fan drive gear system is configured to provide a speed reduction between the fan and the fan drive turbine greater than about 2.5:1.
 13. The gas turbine engine as set forth in claim 12, wherein the gas turbine engine is a high bypass geared aircraft engine having a bypass ratio greater than about six (6).
 14. The gas turbine engine as set forth in claim 13, wherein the bypass ratio is greater than about ten (10).
 15. The gas turbine engine as set forth in claim 13, wherein the gas turbine engine includes a fan pressure ratio of less than about 1.45.
 16. The gas turbine engine as set forth in claim 15, wherein the fan is configured to rotate with a fan tip speed that is less than about 1150 ft/second.
 17. The gas turbine engine as set forth in claim 16, wherein the fan drive turbine is configured to operate at a pressure ratio that is greater than about 5:1.
 18. The gas turbine engine as set forth in claim 17, wherein the second turbine section operates at a higher pressure than the fan drive turbine and is configured to drive the second compressor section.
 19. A method of designing a gas turbine engine comprising: defining a fan configured for rotation about an axis; configuring a compressor section, a combustor section, and a turbine section to define a core flow path; configuring a fan drive gear system for driving the fan at a speed different than the turbine section; defining a lubricant system to include a lubricant pump for delivering lubricant to an outlet line, said outlet defined to split into at least a hot line and into a cool line, said hot line being configured for directing lubricant primarily to locations in the gas turbine engine that are not intended to receive cooler lubricant, and said cool line configured for directing lubricant through one or more heat exchangers for cooling lubricant, and said cool line is configured for routing to the fan drive gear system; and at least one of said one or more heat exchangers is configured as a fuel/oil cooler at which lubricant will be cooled by fuel leading to the combustion section, and said fuel/oil cooler being downstream of a point where said outlet line splits into said at least said hot line and said cool line, such that said hot line is not directed through said fuel/oil cooler.
 20. The method as set forth in claim 19, including configuring said locations in the engine that are not intended to receive cooler lubricant to include bearings associated with at least the turbine section.
 21. The method as set forth in claim 20, wherein said cool line is configured to supply lubricant to a journal bearing in said fan drive gear system.
 22. The method as set forth in claim 19, including defining a valve for selectively supplying lubricant from said hot line to the fan drive gear system when additional lubricant is necessary.
 23. The method as set forth in claim 19, including configuring a valve to be positioned on said cool line, at a location downstream of a point where said hot line splits off from said cool line, and configuring said valve for splitting said cool line into two lines, with a first line being directed through said one or more heat exchanger, and a second line being directed through at least one other cooler.
 24. The method as set forth in claim 19, including defining the system such that lubricant in said hot line does not intermix back into said cool line until a point after said cool line has been routed to a bearing for said fan drive gear system.
 25. The method as set forth in claim 19, including defining the system such that lubricant in the hot line is intermixed with a portion of the lubricant from the cool line prior to being delivered to said locations in the gas turbine engine that are not intended to receive cooler lubricant. 